1. Field of the Invention
The present invention relates to supersonic aircraft, joined-wing aircraft and sonic boom reduction, separately and in combination, for long range supersonic cruise aircraft with reduced sonic boom loudness.
2. Description of Related Art
A number of prior patents and technical reports have individually described technologies for joined-wings, low supersonic drag and sonic boom minimization. Further design studies and actual supersonic aircraft revealed shortcomings that reduced and eliminated the perceived advantages of such designs. Although tandem and connected wing designs have been proposed in prior patents, such designs are unfavorable because the wing downwash on the following wing and pitch stability were not taken into account. Downwash created by one wing reduced the lifting efficiency of any nearby or following wing, requiring a greater angle-of-attack to generate the same lift. This reduction in lifting efficiency increased induced drag and reduced pitch stability. Trailing wings carried very little lift or often a down load to trim an aircraft with the center-of-gravity far enough forward for positive pitch stability. Current artificial stability technology allows the trailing wing lift to be slightly increased, but also increases development cost. Downwash is the main reason why biplane and canard airplanes were abandoned and now serve only in applications not primarily based on efficiency.
Additionally, prior patent designs were generally focused on subsonic applications rather than supersonic applications. Box-wings and tip-connected joined-wings claim vortex drag reductions as their primary improvement. However, the contribution of vortex drag to total drag and the vortex drag benefit of interconnection diminish as speed increases supersonically. Another important requirement of supersonic applications is the need to keep cross-sectional area distributions as low as possible and minimize second derivative changes, also known as area ruling, to reduce wave drag. Blunt and unswept wings, blunt connecting elements and superimposed connections have rapid cross-sectional area changes that produce high wave drag supersonically. Supersonic surfaces and bodies need to be sharp or swept greater than the Mach cone angle and need low thickness-to-chord ratios. The Mach cone angle is defined as the inverse cosine of 1/Mach, and is 60 degrees at Mach 2.
Finally, most prior patent designs that address sonic boom only reduce the sonic boom due to aircraft volume or use energy to alter boom pressures. Unfortunately, reducing the sonic boom due to volume results in a sonic boom due to lift that is typically stronger than the combined lift and volume boom. This counter-intuitive result occurs because the lift is concentrated in a length shorter than the vehicle length and the typical area ruled volume superimposes an expansion where the lift is located, mitigating the lift. As for energy techniques, the energy in the boom pressures is equivalent to the entire propulsion system output that balances them, suggesting impractical power requirements.
To help understand some of the benefits achieved in accordance with the present invention, the relevant shortcomings of prior designs, for the present supersonic application, are described below.
U.S. Pat. No. 1,264,037 to Emmons teaches a biplane aircraft with a sweptback forward wing connected to a sweptforward rear wing also connected to an additional mid-fuselage fin. The wings are connected by vertical struts, which cannot provide joined-wing 20¢: in-plane stiffening or box-wing vortex induced drag reduction.
U.S. Pat. No. 1,453,830 to Coakley teaches multiple connected wings with dihedral for the lower wings and anhedral for the upper wings and vertical struts, creating a stiff structure that carries loads in-plane. However, it is only suitable for low speed flight having its connections superimposed, no wing sweep and an excess of bracing struts and wires.
U.S. Pat. No. 2,406,625 to Oglesby teaches multiple, unswept wings of alternating anhedral and dihedral connected at their tip to a common tube and connected at their root to a fuselage. This is recognized to form a truss arrangement, reducing wing loads.
However, the lack of sweep leads to a long tip-connecting tube that would have significant weight and friction drag. Multiple trailing wings have increasing drag due to increasing downwash.
U.S. Pat. No. 2,461,805 to Barker teaches dual wings and dual tails of differing height and tandem fore/aft position at their root and connected at their tips by an interconnecting structure that blends the forward tip into the aft tip. While this does carry loads in-plane, greatly increasing strength and rigidity, the interconnecting structure creates a large spanwise blockage of the flow at the tips unsuitable for compressible flow regimes (Mach greater than 0.5) and creates redundant, closely-coupled surfaces increasing friction and interference drag.
U.S. Pat. No. 2,567,294 to Geraci teaches a tail-less bi-plane in a box-wing arrangement also with dual propeller box-prop arrangement. While there would be a vortex drag reduction, the wings are close causing high interference (downwash) and the vertical interconnecting structures cannot transfer loads in-plane for less strength and rigidity benefit. Further, both wings are highly sweptback, resulting in little torsional stiffness improvement leading to increased weight to resist flutter and aeroelastic elevon reversal. The interconnecting fins are also superimposed with the wings, creating increased area unsuited for compressible flow regimes.
U.S. Pat. No. 3,834,654 to Miranda teaches a box-wing with three relevant (to this patent) advantages. First, in having a rearwardly swept wing connected at the tip, by a vertical fin, to a forwardly swept wing, increased torsional stiffness eases aeroelastic problems. However, the box-wing is not claimed to resist lift loads in-plane, so it is more similar structurally to dual cantilevered wings. Second, reduced transonic and supersonic wave drag is attributed to the greater length over which the wing volume is spread. As a caveat, the patent notes that the superposition of wing/fin connections could increase shock strength and drag, and therefore, suggests: rounding the corners, adding the volume of a streamlined center body or using boundary layer control through suction, blowing or vortex generators. Third, the vertical separation of the wings theoretically makes possible a 40% induced drag reduction subsonically. Unfortunately, this improvement is greatly reduced with practical longitudinal stability requirements. A stable box-wing as shown requires a download from the aft wing to trim and current artificial stability systems would only allow a small upload, so trim drag and flight control system impacts must be considered when assessing benefits.
U.S. Pat. No. 3,942,747 to Wolkovitch teaches a canard swept back and up connected at its tip to an unswept wing, both canard and wing attached to a fuselage. The design is suited to a low speed, ultralight-type aircraft stressing simple construction and rugged design at the expense of performance. The design is in particular claimed suited to flexible lifting portions made material such as canvas.
U.S. Pat. No. 3,981,460 to Ratony teaches tip-joined, counter-swept wings with the upper wing being free from fuselage and vertical fin bracing, increasing load and wing weight. The acute angle and wing superposition at the tip results in high interference and wave drag in compressible flow regimes. Further, the short distance between the control surfaces and the center-of-gravity results in a large trimmed loss of flap lift and increase in drag. Finally, the preferred embodiment describes that elimination of the wing tip vortex eliminates sonic boom. Current theory attributes the formation of sonic boom to the (downward momentum imparted to the air to generate) lift and the volume disturbance. As this design shows no area ruling and a relatively short vehicle length, the theory would indicate greater than average sonic boom at the ground.
U.S. Pat. No. 3,985,317 to Geraci and DeLouise teaches a compact, unswept, staggered box-wing aircraft not suitable for high speed flight.
U.S. Pat. No. 4,053,125 to Ratony teaches a sweptback forward wing joined at the tip to a sweptforward aft wing, both wings mounted on the fuselage with an aft T-tail. The wings do form an interconnected structure capable of carrying loads in-plane for lighter weight. The wings are superimposed at their tips and drawings indicate an attachment structure that reduces channel closure but add further tip volume, causing increased wave drag.
U.S. Pat. No. 4,090,681 to Zimmer teaches both sweptback upper and lower wings of opposing dihedral, joined at their tips with an aft tail. The upper wing is swept more so that its trailing edge is positioned approximately over the leading edge of the lower wing. This arrangement is well illustrated as being able to carry loads in-plane for lighter weight wing structure. However, the tips are connected through a blunt leading edge plate that would have high stress. In addition above about Mach 1.4, the stagger shown between the wings would cause the growth of the forward wing volume to start aligning with the aft wing volume growth, greatly increasing wave drag.
U.S. Pat. No. 4,146,199 to Wenzel teaches a low mounted sweptback wing connected at the tips to a high mounted sweptforward wing extending from either side of a lifting body fuselage. The design is not suited for high speeds or altitudes requiring pressurization with a rectangular fuselage cross-section. As the lifting body is very low aspect ratio with very short end plates, induced drag would be very high for the lift developed on the fuselage. The wing tip connections are staggered and separated by a blending structure, reducing interference drag but resulting in high stresses in the blending structure.
U.S. Pat. No. 4,365,773 to Wolkovitch teaches a wing connected to an aft fin and sweptforward connecting its tip, through a streamlined element, to a sweptback wing in the elevational plane of the wing. The sweptback wing is attached to the fuselage at a lower elevation than the sweptforward wing. The higher, fin-mounted wing opens up the angle at the wings"" connections. All drawings show high aspect-ratio (span divided by chord) wings for subsonic applications. Large streamlined members are shown for staggered wing connections. These lighten the weight for load transfer but incur a friction area increase. The streamlined members are also shown with their maximum thickness aligned with the sweptback wing maximum thickness, and fin-wing connections are superimposed. Both superpositions reduce the drag divergence Mach number and increase transonic and supersonic wave drag. The claims further teach that the structural box should be strengthened in opposite corners whose diagonal is more out-of-plane with the connecting wing. However, the structure of the stiffening would need to be different for multi-spar wings with mid-span connections.
U.S. Pat. No. 4,541,593 to Cabrol teaches upper and lower wings connected at the tips and hingedly secured to the fuselage. The hinge connection reduces the wing""s resistance to compressive buckling, resulting in an increase in wing weight relative to a stiff connection. The wings"" lift is also maximized at mid-span locations, which increases stress and wing weight.
U.S. Design Pat. No. 292,203 to Legeti shows an aircraft very similar to U.S. Pat No. 3,942,747 to Wolkovitch except for a box-wing interconnection and ducted pusher prop. The design is not suited for flight in compressible flow regimes.
U.S. Design Pat. No. 292,911 to Argondezzi shows a box-wing arrangement similar to U.S. Pat. No. 3,834,654 to Miranda except for an additional unswept wing also attached to the vertical interconnecting fins between the upper and lower wings. The design is not suited for flight in compressible flow regimes.
U.S. Pat. No. 4,856,736 to Adkins, McDonald and Sivertsen teaches a front, sweptback wing mounted high and an aft sweptforward wing mounted low on the fuselage like patent 4,053,125 to Ratony except that the wings are joined at the tip in tandem with an interconnecting structure to link the wing structures. While reducing interference drag of a superimposed wing tip connection, the instability and inability to efficiently lift from the aft wing and trim the aircraft eliminate much of the claimed advantages.
U.S. Design Pat. No. 304,821 to Ratony shows an arrangement similar to design U.S. Pat. No. 292,911 to Argondezzi except that the wing tips are joined in tandem to a common streamlined member with triple winglets. The unswept center wing and pusher props are not suited for flight in compressible flow regimes.
U.S. Pat. No. 5,046,684 to Wolkovitch teaches an aircraft like his U.S. Pat. No. 4,365,773 with the addition of pivoting propulsion devices mounted and braced at the wing tip connection. It teaches improvements for short and vertical takeoff and landing designs for subsonic cruise speeds.
U.S. Pat. No. 5,503,352 to Eger teaches a low fineness ratio, box-wing aircraft with a ducted pusher prop not suited for flight in compressible flow regimes.
U.S. Pat. No. 5,615,846 to Shmoldas, Hutchings and Barlow teaches a planer joined-wing with a mid-span connection as a means of actuating an extendable wing through an umbrella-like action. The planer wings cannot provide bracing of lift loads.
U.S. Pat. No. 5,750,984 to Morgenstern teaches a supersonic aircraft with a moveable control surface at the nose and a sonic boom minimized area/lift distribution. The control surface is used to regulate the strength of the boom minimized, triangular, nose pressure spike to match non-standard conditions or to eliminate the spike drag penalty during over water operation where somewhat louder boom has been accepted.
U.S. Pat. No. 5,899,409 to Frediani teaches a box-wing aircraft like U.S. Pat. No. 3,834,654 to Miranda with the addition of a greater than 400,000 kg maximum take-off weight and fuselage with multiple passenger decks upon which both wings are mounted. The smaller vertical separation of the fuselage-mounted wings reduces the induced drag benefits and amplifies previous comments regarding stability reduction or increased trim drag.
U.S. Pat. No. 5,899,410 to Garrett teaches coplaner joined wings. Since the wings are coplaner, structural bracing of lift load is not possible, and therefore, is carried like cantilevered wings. Joints show coplaner fairings that increase area at connections increasing transonic and supersonic drag.
U.S. Design Pat. No. 417,184 to Hartmann et al. discloses an ornamental design for a supersonic business jet.
In addition to the above-noted patent literature, joined-wing optimization studies have been performed for a transonic (Machxcx9c0.8) transport aircraft application. These studies focused on configuring a joined-wing for optimum aerodynamic and structural trade-offs, and determined that the weight savings of inboard joint locations outweighed the aerodynamic advantages of tip joint locations. They also found that a download on the aft wing or tail was needed to trim the aircraft 20 percent statically stable. They further determined that their optimized joined-wing actually performed worse than the conventional design for this application. However, applications with a long wing chord, short distance between wing and tail, very low wing t/c and no maximum lift (CLmax) limiting would be more favorable for joined-wing aircraft. Supercruise applications typically have all of these characteristics.
High supersonic L/D technology also is documented in technical literature. The literature describes L/D improvements due to the area-rule, optimum lift distribution including nacelle interference, wing leading edge sweep-back greater than the Mach cone angle, high trailing edge notch ratio (longitudinal distance the wing trailing edge root is ahead of the tip versus total wing length), and podded nacelles mounted under the wing behind the wing maximum thickness. Unfortunately, later more in-depth NASA SST and HSCT programs proved these wing and nacelle features to be structurally heavy and even infeasible. The high wing sweeps and resulting long structural load paths are poor at resisting aeroelastics and flutter. Moreover, propulsion nacelles cantilevered behind the maximum thickness increase flexure. NASA programs were unable to get net aero/structural improvements from positive notch ratio and even had to minimize the portion of leading edge sweep-back greater than the Mach cone angle to improve subsonic performance. The inability to take advantage of these high L/D technologies was due to the low stiffness of a conventional cantilevered wing.
Since the nacelle is generally the most concentrated source of pressure drag, recovering its drag on the other surfaces of the vehicle has been proposed by designers. The nacelle pressure drag may be recovered by altering the wing camber slope. Previous to this technique, flow calculations without nacelles were used to determine the wing camber slopes, or flexure, for a minimum drag, lift distribution. It was noticed that when nacelle pressures impinged on the wing, the lift distribution changed, and it was speculated that the wing should be, at least partially, reflexed to restore the original minimum drag, lift distribution. Nacelle pressure disturbance impingement calculations were made using an axisymmetric equivalent body and linearized theory. When the angle of the wing camber slopes was changed (reflexed) to restore the original lift distribution, the drag was improved, but the drag was reduced the most when only one-half to two-thirds of the slope change was used. In wind tunnel tests, recovering the nacelle pressure drag through wing reflexing often resulted in half or less of the improvement expected due to additional inaccuracies in calculation and manufacturing resolution. Wing reflexing diminished in importance as a design technique due to these marginally effective results.
Shaped sonic boom, as the term is used herein, specifically refers to altering the source pressure disturbance such that a non N-wave shape is achieved at the ground. Shaping sonic boom is capable of loudness reductions of 15-20 dB or higher, reductions with no additional energy requirement (beyond that already needed for flight), and successful demonstrations beyond the near-field in wind tunnel tests. The key to understanding shaped sonic boom, using shaping to minimize loudness and applying the concept to practical aircraft designs, starts with understanding how aircraft pressure disturbances change as they propagate to the ground.
Shaped sonic booms are only achieved deliberately. No existing aircraft creates a shaped sonic boom that persists more than a fraction of the distance to the ground while flying at an efficient cruise altitude. Typical source pressure disturbances quickly coalesce into an N-wave, a shape with the largest shock magnitudes possible from a given disturbance. Coalescence occurs because the speed at which the pressure disturbance propagates varies slightly, proportional to its pressure and thereby temperature (speed of sound=sqrt (xcex3RT). Since the front of a supersonic aircraft generates an increase in ambient pressure, and the rear generates a decrease in pressure, the variation in propagation speed causes aircraft pressure disturbances to stretch-out as they propagate to the ground. As the disturbances stretch-out they also tend to coalesce because shocks travel at halfway between the speed of the lower pressure ahead and higher pressure behind them. In general, to keep pressures from coalescing the pressures at the nose must have a large compression and the pressures at the tail must have an expansion, with the pressures in between constrained to weak compressions and expansions. This causes the ends of the signature to stretch out faster than the pressures between them, resulting in non N-wave sonic boom at the ground.
Acoustic studies have shown that sonic boom loudness at audible frequencies correlates with annoyance. Meaning that to achieve acceptable supersonic flight over land, the loudness of a sonic boom needs to be minimized. The audible frequencies in a sonic boom occur in the rapid pressure changes, or shocks, at the beginning and end of the typical xe2x80x9cNxe2x80x9d waveform. Shocks become quieter with decreasing magnitude and with increasing rise time (of the pressure change). However, the shock rise time is inversely proportional to its magnitude (although there is a large variability around this relationship in measurements). Therefore, the audible sound pressure goes down dramatically with decreasing shock magnitude. For example, two shocks of half the pressure of a single shock are about 6 dB quieter, and one shock of half the pressure is about 9 dB quieter. In summary, minimizing shock magnitude minimizes loudness.
The lowest shock magnitude from a shaped sonic boom may be determined theoretically. Sonic boom minimization methodology calculates the minimum shock strength possible subject to a compression slope input (fraction below minimum coalescence slope, 0 for N-wave, 1 for flat-top signature) for a given vehicle length and weight at the desired flight conditions. This is the lowest shock magnitude possible, assuming that any compression between the shocks must have a constant, linear slope. Any shaped sonic booms not corresponding to a minimized source pressure distribution, like those with multiple front or rear ground shocks, is basically using the same ideas (weak enough pressures between the ends to prevent shock coalescence) with a non-optimally minimized distribution.
The minimized source pressure distribution (non-dimensionalized as an f-function) and the corresponding equivalent area distribution may be calculated. Equivalent area is a term that is defined as the axisymmetric Mach angle-cut area distribution needed to generate a corresponding pressure distribution. A given lift distribution can be converted into a corresponding equivalent area distribution. Equivalent area provides a more geometric way to examine the design consequences of a sonic boom minimized pressure distribution. A shortcoming of sonic boom minimization attempts is that they only consider minimization directly below the aircraft, instead of a more general approach of how to simultaneously reduce sonic boom off to either side and directly under the flight path.
The present invention has been developed in view of the foregoing, and to address other deficiencies of the prior art.
The present invention provides improvements to joined-wing aircraft and sonic boom minimization that are necessary for a supersonic application.
An embodiment of the present invention provides a tail-braced wing aircraft capable of long supersonic cruise range.
Another embodiment of the present invention provides sonic boom minimization, which is a technology for altering the shape of the boom waveform to minimize audible higher frequency sounds. This is achieved by tailoring the area and lift distribution. Loudness reductions of 30 times (audible pressure reduction or 15 dBA) or higher may be achieved.
Relative to the prior designs noted above, the present designs may reduce structural weight of joined-wings, reduce wave drag from the joints, reduce supersonic induced drag through the use of supercruise optimized lifting surface shape and load distribution and/or improve the shape of the minimized boom nose pressure spike and nose shape without regulating the strength of the spike.
An aspect of the present invention is to provide an aircraft comprising a fuselage, a swept-back wing having a wing root connected to the fuselage, a fin mounted on an upper surface of the fuselage at a rear portion of the aircraft, and a bracing tail. The bracing tail has a bracing tail tip connected to the wing inboard of a tip of the wing, and a bracing tail root connected to the fin at a higher point vertically than the connection between the wing root and the fuselage. The connection between the wing and the bracing tail tip comprises means for maintaining a substantially constant area distribution, i.e., the area distribution is more constant than the distribution for a separate wing and bracing tail.
Another aspect of the present invention is to provide an aircraft comprising a fuselage, a swept-back wing having a wing root connected to the fuselage, a fin having a fin root mounted on an upper surface of the fuselage at a rear portion of the aircraft, and a bracing tail having a bracing tail tip connected inboard of a tip of the wing at a point that is greater than 50 percent of the distance from the wing root to the wing tip, and a bracing tail root connected to the fin at a point that is greater than 20 percent of the distance from the fin root to a tip of the fin.
A further aspect of the present invention is to provide a supersonic cruise aircraft comprising lift and trim surfaces, and a propulsion system capable of generating a non-axisymmetric supersonic pressure disturbance that impacts the lift and/or trim surfaces, wherein the lift and/or trim surfaces define a camber line having a slope selected to cancel about one half of the supersonic pressure disturbance.
Another aspect of the present invention is to provide a supersonic aircraft comprising a fuselage, a swept-back wing connected to the fuselage, a fin connected to the fuselage, and a bracing tail connected to the wing and fin, wherein the fuselage comprises a blunt nose having a camber line that slopes upward toward the front of the aircraft.
A further aspect of the present invention is to provide a supersonic cruise aircraft comprising a body, and a lifting surface mounted on the body, wherein the aircraft has a combined lift/area that produces a shaped sonic boom minimized distribution.
These and other aspects of the present invention will be more apparent from the following description.